1. Field of the Invention
This invention relates generally to composite material structures and, more specifically, to methods for manufacturing composite material structures. In one particular embodiment, this invention relates to single cure composite material structures formed with one or more separable and replaceable portions using resin transfer molding methods.
2. Background
Composite materials, such as carbon fiber present in an organic matrix, have been used to produce corrosion resistant and light weight structures. These structures typically weigh about 25% less than structures made of lightweight metals, such as aluminum, while at the same time offering similar strength to these metals. Composite structures have offered performance improvements in terms of lower weight, increased stiffness, and long fatigue life, and have been used to fabricate a wide variety of structures including, for example, aircraft structures (such as fuselage shell components, wing sections, tail sections, etc.). These composite structures have typically been manufactured by time consuming methods, such as hand placement and autoclave cure. For example, in one process for manufacturing hollow landing gear doors, resin impregnated prepreg materials have been hand laid-up on mandrels with a non-stick dividing material placed between sections of the prepreg material so as to isolate one wall of the landing gear door structure from the remaining walls of the structure during lay-up and cure, thus forming a removable wall section through which the mandrels could be later removed, and then the wall section replaced to form the finished door. However, traditional hand lay-up autoclave composite construction is labor intensive, uses high cost resin-impregnated raw materials that require special storage, and therefore often results in costly finished structures.
Many composite structures, such as aircraft wings and tails, are manufactured in multiple and separate parts so as to provide needed access to the interior of the structure for insertion of parts prior to final assembly. For example, traditional composite structure methods of producing a complex aircraft part such as a wing typically involve separate fabrication and curing of skeleton and skin components, and involve assembly after all internal parts are installed. Not only is this traditional method labor intensive, but the skins typically do not exactly match the skeleton, and customizing measures are usually needed during assembly to correct the gaps or interferences between the skins and the skeleton, requiring a substantial amount of time to achieve a uniform and consistent fit.
Single cure closed tool Resin Transfer Molding (xe2x80x9cRTMxe2x80x9d) is a method of forming composite structures in which one piece composite parts are cured under high temperature and high pressure in a closed tool using a single cure cycle. In a conventional RTM procedure, dry carbon fiber fabrics are applied to a mandrel, loaded into a tool, then injected with resin and cured to form a completed structure, thus eliminating the need for assembly of the traditional individually fabricated parts such spars and skins. For example, using one single cure RTM method in the manufacture of wing flaps, dry carbon fiber braided material is fitted over or wrapped around inner mandrels and placed inside a mold. The mold is closed around the mandrels and placed in a heated press. A vacuum is then drawn on the interior of the mold and a predetermined volume of heated resin injected into the mold to saturate the dry braided material. A relatively high pressure is applied to the resin, and the tool is heated to cure temperature. After the cure cycle is complete, the mold is removed from the press, and the mandrels and composite part are removed from the mold to cool, prior to removal of the mandrel from one of the open ends of the composite part.
Single cure closed tool RTM methods produce structures which have a relatively low labor content, and because dry fiber material is used no special storage of resin impregnated fiber material prepreg is required. Because a one piece composite structure is produced in a single step, considerable labor cost is saved by eliminating assembly steps. This is in part because of the ability to mold completed assemblies and also because parts may be cured in the molding tool, instead of requiring that the parts in the molding tool be move to an autoclave for curing. Use of a closed tool RTM method enables a closed structural section (e.g., box-like structures) to be fabricated in one piece, eliminating the assembly steps necessary when the spars and skins are fabricated separately. The tools for RTM curing typically allow for the tool itself to be heated and allow for pressure to be applied hydraulically to the resin whereas traditional methods of composite fabrication rely on the autoclave to apply heat and pressure.
Single cure closed tool RTM processes have been used to produce simple aircraft composite structures (e.g., wing flaps, drain mast) more quickly and efficiently than other methods. RTM processes have been tried using combinations of liquid injected resin and prepreg materials, and a combination of injected liquid resin and prepreg materials has been employed to produce a landing gear door. However, one factor that has hampered the use of closed tool RTM processes to produce larger and more complex hollow aircraft structures (e.g., wings and tails) is lack of access to the inside of the-completed composite structure. As previously described, access is typically required for such structures in order to install supporting structures such as ribs, and equipment such as fuel pumps and gages. For example, ribs may be desirable in an aircraft wing fuel tank to prevent the fuel from xe2x80x9csloshingxe2x80x9d out towards the wing tip when the airplane rolls into a one wing down attitude. However, the current technology single cure cycle RTM part has no accessibility except from each end.
Disclosed herein are single cure RTM processes that may be employed to produce high quality composite structures having relatively lightweight and relatively low fabrication costs, and which offer the advantages of a single in-tool cure cycle while at the same time allowing internal equipment and/or internal structure to be added to a post cure structure. In one embodiment are methods that advantageously allow the use of single cure closed tool RTM process to fabricate relatively large and complex closed section structures such as an aircraft wing or tail fin (e.g., box like structures with a front spar, a rear spar, and with top and bottom skins) in one piece, while permitting full access to virtually any pre-determined part of the internal structure, thus eliminating extra assembly steps necessary when the spars and skins for such components are fabricated separately.
Advantageously, using the disclosed RTM process a relatively low cost one piece structure (e.g., wing or tail fin section) may be fabricated in a single cure cycle, but in a manner that allows access to the inside of the structure after cure (e.g., to add internal ribs to control fuel slosh/pressure and to add equipment such as fuel pumps and gauges). Surprisingly, the disclosed RTM process may be used to produce a composite structure having a separable part (e.g., separable top skin, separable bottom skin, or any other portion of the structure thereof) that is substantially separable from the rest of the structure (e.g., the skeleton) after cure. In one embodiment, the separable part (or any portion thereof) may be easily reconnected to the structure by virtue of the fact that it is molded as a matching part of the structure during the cure cycle.
When the disclosed single cure RTM processes are used to manufacture an aircraft wing or tail structure having a separable skin portion, the separable skin and the skeleton are cured in the same tool and with the same cure cycle. Therefore, the disclosed processes offer a major benefit over traditional processes that rely on fabrication of individually fabricated skins and associated skeleton of an aircraft wing or tail because a separable skin and the adjacent skeleton substructure share a substantially perfectly matching surface due to the fact that they are cured together. These parts are therefore substantially perfectly matched at their common surfaces even after separation of the skin portion from the skeleton substructure, virtually eliminating the issues and problems associated with fitting together and reconnecting the separated skin to a skeleton substructure.
Advantageously, the separable skin portion (e.g., of an aircraft wing or tail) may be separated after cure and removal of the entire structure from the molding tool and may remain separated from the remainder of the structure until all installations, sealing and inspections have been completed. Further advantageously, final assembly of the structure typically does not require elaborate tooling to hold and align the separable skin with the skeleton substructure, as the separable skin portion may only be oriented in one way relative to the skeleton so as to achieve a substantially perfectly matching fit with the skeleton. Thus, in one embodiment, wings for fighters, trainers, or business jets, as well as tails for larger airliners may be cured in a single cure cycle in a mold and still allow access to the internal structure due to the provision for separation of the top or bottom skin of such parts from the substructure and easy reinstallation of same to the substructure.
Further advantageously, the disclosed single cure RTM processes may be employed to fabricate structures having sandwich cored skins, such as a honeycomb core surrounded by layers of carbon fiber. Composite skins formed in this manner offer lightweight strength and resistance to buckling due to presence of, for example, a honeycomb core, but at the same time may be advantageously fabricated using an efficient and relatively low cost RTM process that allows fabrication of such a skin without resin-saturation of the honeycomb core, as would occur with conventional RTM processes. In this regard, the honeycomb core may be surrounded by prepreg fiber material that prevents contact of the honeycomb core with the injected resin, and which includes resin advantageously selected to have the same curing characteristics as the RTM resin, thus allowing the entire structure to cure during the same RTM curing cycle.
Thus in one respect disclosed herein is a composite structure, and a method of forming the same which includes: positioning first and second material pieces adjacent each other within a mold cavity and with a separating film disposed between the first and second material pieces, the first material piece including an impregnable fiber material piece; introducing a first resin into the mold cavity to contact and substantially impregnate the first material piece with the first resin; allowing the first resin that substantially impregnates the first material piece to cure within the mold cavity to form a first part of the composite structure; and in which the separating film is substantially impermeable to the first resin. The separating film may be selected to substantially prevent the first resin from contacting the second material piece and so that the first resin substantially does not adhere to the separating film during the curing of the first resin so that the separating film may be substantially separable from each of the composite structure first part and the second material pieces after the first resin is cured.
The separating film may have opposing first and second surfaces; and the method may include positioning at least one surface of the first material piece in contact with the first surface of the separating film and positioning at least one surface of the second material piece in contact with the second surface of the separating film; so that the separating film substantially prevents the first resin from contacting the at least one surface of the second material piece and so that the composite structure first part and the second material pieces are substantially separable from each other after the first resin is cured.
In one embodiment, a first material piece may include one or more material pieces with at least one surface of the first material piece having an open area defined therein between a first area of the first material piece and a second area of the first material piece; and wherein the positioning includes positioning the at least one surface of the first material piece adjacent a first surface of the separating film with a resin transport film disposed therebetween so that the resin transport film may be in a position effective to transport the introduced resin from the first area of the first material piece across the open area to the second area of the first material piece. In another embodiment, the method may further including forming an open area within a surface of the composite structure first part after separating the first and second composite structure parts; wherein the open area may be formed so as to be covered by the second composite structure part when the first and second composite structure parts are refitted together.
In another respect, disclosed herein is a cured composite structure, and a method of forming the same which includes: positioning a first material piece at least partially around an outer surface of at least one mandrel, the first material piece including a dry fiber material piece; positioning a second material piece adjacent the first material piece with a separating film disposed therebetween to form an uncured composite structure; positioning the uncured composite structure and the mandrel within a closed mold cavity; introducing a first resin into the closed mold cavity to contact and at least partially impregnate the first material piece with the first resin; allowing the first resin that at least partially impregnates the first material piece to cure within the mold cavity to form at least a portion of a cured composite substructure, the cured composite substructure and the second material piece together forming the cured composite structure; removing the cured composite structure from the closed mold cavity; and removing the mandrel from the cured composite substructure to form a hollow cured composite structure. In this method, the separating film may be selected to be substantially impermeable to the first resin, and the cured composite substructure may be substantially separable from the second material piece after removal from the closed mold cavity.
In another respect, disclosed herein is a hollow cured composite aircraft structure, and a method of forming the same using a single cure RTM process which includes: positioning a first fiber material piece at least partially around an outer surface of at least one mandrel, the first material piece including a dry fiber material piece; positioning a second fiber material piece adjacent the first fiber material piece with a separating film disposed therebetween to form an uncured composite aircraft structure, wherein the separating film may have opposing first and second surfaces, and wherein at least one surface of the first material piece may be positioned in contact with the first surface of the separating film and at least one surface of the second material piece may be positioned in contact with the second surface of the separating film; positioning the uncured composite aircraft structure and the mandrel within a closed mold cavity; introducing a first resin into the closed mold cavity to contact and at least partially impregnate the first fiber material piece with the first resin; allowing the first resin that at least partially impregnates the first fiber material piece to cure within the mold cavity during a single cure cycle to form at least a portion of a cured composite aircraft substructure; allowing a second resin that at least partially impregnates the second fiber material piece to cure within the mold cavity during the single cure cycle to form at least a portion of a cured composite aircraft structure skin, the cured composite aircraft substructure and the cured composite aircraft structure skin together forming a cured composite aircraft structure; removing the cured composite aircraft structure from the closed mold cavity; and removing the mandrel from the cured composite substructure to form the hollow cured composite aircraft structure.
In this method, the separating film may be selected to be substantially impermeable to the first and second resins so that the separating film substantially prevents the first resin from contacting the at least one surface of the second fiber material piece that may be in contact with the second surface of the separating film, and so that the separating film substantially prevents the second resin from contacting the at least one surface of the first fiber material piece that may be in contact with the first surface of the separating film. In the practice of method, the materials may be selected so that the first resin substantially does not adhere to the separating film during the curing of the first resin and the second resin substantially does not adhere to the separating film during the curing of the second resin so that the cured composite aircraft substructure is substantially separable from the cured composite structure skin and so that the separating film is substantially separable from each of the cured composite aircraft substructure and the cured composite aircraft structure skin after the first and second resins are cured and after removal of the cured composite aircraft structure from the closed mold cavity. The method may further include separating the separating film from the cured composite aircraft substructure and the cured composite aircraft structure skin after removal of the composite aircraft structure from the mold cavity, and then refitting together the hollow cured composite aircraft substructure and the cured composite aircraft structure skin.